This invention relates to a low pressure nuclear thermal rocket (LPNTR), and more particularly, an LPNTR that utilizes radial outflow of propellant through radial conical fuel assemblies in a shell-type (spherical, cylindrical, or a combination of both) reactor.
Two decades ago the Space Nuclear Propulsion Office (SNPO) provided guidelines for the design of the Nuclear Engine for Rocket Vehicle Application (NERVA) flight engine then under development. These guidelines provided that reliability and the achievement of the highest probability of mission success were the major design criteria. Next in the order of importance was performance as measured in terms of specific impulse. Then, the engine design should attempt to keep the overall weight as low as possible within the bounds allowed by funds available for development. The relative priorities indicated are equally applicable today to the Space Exploration Initiative recently undertaken by NASA and the Department of Energy (DOE). These overall performance priorities: reliability; specific impulse; and, engine thrust/weight ratio; are basic to the selection of the LPNTR concept.
The relative importance of specific impulse (I.sub.sp) versus thrust/weight (T/W) ratio is a matter of considerable importance to the optimization of a solid core nuclear thermal rocket (NTR) for space flight. High engine T/W ratio is relatively important to the use of an NTR as a second stage of a launch vehicle or for some military space applications. However, compared to I.sub.sp, high engine T/W ratio is relatively unimportant for Mars missions and most other missions of potential interest to NASA.
The two engine systems (NRX/EST and XE) tested in the NERVA program utilized a hot-bleed engine cycle with a single turbopump assembly (TPA). A reliability assessment report for the flight version of such an engine noted that the TPA failure rate represented more than fifty percent of the total engine unreliability. Subsequently the NERVA flight engine design was redirected to the use of dual TPA's in the full-flow engine cycle. To achieve the desired engine system reliability, redundancy of major valves was added. The resulting engine system concept that existed at the termination of the NERVA program was quite complex.
The goal of the LPNTR is an engine system without propellant feed pumps: the propellant run tank pressure alone forces the propellant through the engine. An early study based on then current NERVA fuel materials technology indicated potential feasibility for such engines with chamber pressures of about 20 psi and propellant tank pressures of less than 50 psi. Advanced fuel materials technology coupled with an innovative reactor configuration allows extension to even lower chamber pressures and the possibility of greatly improved specific impulse. LPNTR offers attractive potential for achieving a relatively high rating with respect to reliability, specific impulse, and engine thrust/weight ratio.
Accordingly, it is an object of the present invention to provide a nuclear thermal rocket which has high reliability, and high specific impulse with a compatible engine thrust/weight ratio.
Another object of the present invention is to provide a nuclear thermal rocket which is simple in design and operation so as to provide the very high reliability appropriate to manned space flight.
Yet another object of the present invention is to provide a low pressure nuclear thermal rocket that utilizes radial outflow of propellant through radial conical fuel assemblies in a shell-type (spherical, cylindrical, or a combination of both) reactor.